This invention relates generally to the field of temperature resistant components and in particular, to an article comprising a ceramic matrix composite component of a gas turbine engine such as a combustion liner adapted for backside radiative cooling.
Gas turbine engines are known to include a compressor section for supplying a flow of compressed combustion air, a combustor section for burning fuel in the compressed combustion air, and a turbine section for extracting thermal energy from the combustion air and converting that energy into mechanical energy In the form of a rotating shaft. Many components that form the combustor and turbine sections are directly exposed to hot combustion gasses, for example, the combustor liner, the transition duct between the combustor and turbine sections, and the turbine stationary vanes and rotating blades and surrounding ring segments.
It is also known that increasing the firing temperature of the combustion gas can increase the power and efficiency of the combustion turbine. Modem high efficiency combustion turbines have firing temperatures that exceed temperatures of about 1,600xc2x0 C., and even higher firing temperatures are expected as the demand for even more efficient engines continues. Thus, the cobalt and nickel based superalloy materials traditionally used to fabricate the structural gas turbine components must be aggressively cooled and/or insulated from the hot gas flow in order to survive long term operation in the aggressive high temperature combustion environment.
Ceramic matrix composite (CMC) materials have many potential applications in high temperature environments due to their ability to withstand and operate at temperatures in excess of those allowed for a non-insulated superalloy component. However, oxide and non-oxide CMC""s characteristically can survive temperatures in excess of about 1,200xc2x0 C. for only limited time periods in a combustion environment. Furthermore, oxide-based CMC""s cannot be cooled effectively with active cooling systems due to their low thermal conductivity and their limitations in cooling fluid path design due to manufacturing constraints. For example, convective cooling requires a precise amount of cooling air to flow across the backside of the component; however, although this can be done with appreciable difficulty by controlling air flow splits via sizing cooling air exit holes, undesirable air leakage can often result which is difficult to inhibit if not prevent. For another example, convective cooling causes continuous airflow that can cause thermal shock conditions. Moreover, although non-oxide based CMCs can be aggressively cooled to withstand temperatures above about 1200xc2x0 C., they are subject to environmental degradation that limits their useful life. Thus, to increase the operating temperature range and useful life for CMC components, a high temperature ceramic insulation can be used. However, use of such insulation to cover the CMC substrate does not resolve need to cool the CMC substrate.
Accordingly, there is a need to enhance the operation of CMC structural gas turbine components in a high temperature environment. There is also a need to cool a CMC substrate in an easy and inexpensive manner.
Accordingly, an enhanced operation of CMC structural gas turbine components in a high temperature environment is provided. An easy and inexpensive method of cooling a CMC substrate is also provided, as well an accompanying article of manufacture.
One aspect of the present invention involves an article of manufacture, comprising a ceramic matrix composite composition having a frontside and a backside, and coated with a ceramic insulating material on the frontside and coated with a high temperature emissive material on the backside; and a metal element spaced apart from the ceramic matrix composition and defining a gap between the metal element and the ceramic matrix composite, whereby at least a portion of thermal energy exposed to the ceramic insulating material is emitted from the high temperature emissive material to the metal element.
Another aspect of the present invention involves a component of a gas turbine engine, comprising: a ceramic matrix composite composition having a frontside and a backside, and coated with a ceramic insulating material on the frontside and coated with a temperature emissive coating on the backside; and a conductive metal element spaced apart from the ceramic matrix composition and defining an air gap between the metal element and ceramic matrix composite.
Another aspect of the present invention involves a combustion liner adapted for use in a combustion turbine engine, comprising: a ceramic matrix composite composition having a frontside and a backside, and coated with a ceramic insulating material on the frontside and coated with a temperature emissive coating on the backside; a metal element spaced apart from the ceramic matrix composition and defining an air gap between the metal material and ceramic matrix composite; and a platform adapted to secure the ceramic matrix composition relative to the metal element and to maintain the air gap.